摘要 :
The primary objective of the HIFiRE-1 flight experiment was to investigate boundary layer transition at high Mach numbers. The flight geometry was a 7-degree half-angle axisymmetric cone with a nose bluntness of 2.5 mm radius. An ...
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The primary objective of the HIFiRE-1 flight experiment was to investigate boundary layer transition at high Mach numbers. The flight geometry was a 7-degree half-angle axisymmetric cone with a nose bluntness of 2.5 mm radius. An anomaly in the exoatmospheric pointing manoeuvre resulted in much higher re-entry angles of attack during transition than initially anticipated, of approximately 5° -13°. Pre-flight ground experiments were conducted at angles of attack below 6°. This paper presents boundary layer instability measurements obtained in the Oxford High Density Tunnel (HDT) at the Reynolds number, Mach number, and angles of attack as experienced during the re-entry transition phase of flight. A sweep of Reynolds number and angles of attack were completed, to cover the range of conditions experienced during the HIFiRE-1 transition process. High-frequency surface pressure fluctuation results are given at a relatively fine resolution of azimuthal angles between the windward and leeward ray. Particular focus was given to the region where an indented transition front was observed in flight (φ ≈ 20° - 60°). The HDT results showed no evidence of this indented front, as the boundary layer destabilized directly when moving from the windward to the leeward ray. The windward ray boundary layer was more stable in the HDT than in flight. For azimuths (φ = 45° to the leeward meridian, the HDT transition occured at lower Re_x than in flight. These results are at a relatively high wall-to-total temperature ratio (≈ 0.64 instead of ≈ 0.18 for flight) and provide further insight into the three-dimensional laminar-to-turbulent transition experienced in flight.
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摘要 :
The primary objective of the HIFiRE-1 flight experiment was to investigate boundary layer transition at high Mach numbers. The flight geometry was a 7-degree half-angle axisymmetric cone with a nose bluntness of 2.5 mm radius. An ...
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The primary objective of the HIFiRE-1 flight experiment was to investigate boundary layer transition at high Mach numbers. The flight geometry was a 7-degree half-angle axisymmetric cone with a nose bluntness of 2.5 mm radius. An anomaly in the exoatmospheric pointing manoeuvre resulted in much higher re-entry angles of attack during transition than initially anticipated, of approximately 5° -13°. Pre-flight ground experiments were conducted at angles of attack below 6°. This paper presents boundary layer instability measurements obtained in the Oxford High Density Tunnel (HDT) at the Reynolds number, Mach number, and angles of attack as experienced during the re-entry transition phase of flight. A sweep of Reynolds number and angles of attack were completed, to cover the range of conditions experienced during the HIFiRE-1 transition process. High-frequency surface pressure fluctuation results are given at a relatively fine resolution of azimuthal angles between the windward and leeward ray. Particular focus was given to the region where an indented transition front was observed in flight (φ ≈ 20° - 60°). The HDT results showed no evidence of this indented front, as the boundary layer destabilized directly when moving from the windward to the leeward ray. The windward ray boundary layer was more stable in the HDT than in flight. For azimuths (φ = 45° to the leeward meridian, the HDT transition occured at lower Re_x than in flight. These results are at a relatively high wall-to-total temperature ratio (≈ 0.64 instead of ≈ 0.18 for flight) and provide further insight into the three-dimensional laminar-to-turbulent transition experienced in flight.
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This paper presents results of a system study of transpiration cooled thermal protection systems for Earth re-entry. The cooling performance for sustained hypersonic flight and transient re-entry of a blunt cone geometry is assess...
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This paper presents results of a system study of transpiration cooled thermal protection systems for Earth re-entry. The cooling performance for sustained hypersonic flight and transient re-entry of a blunt cone geometry is assessed. A numerical model is developed which employs semi-empirical aerodynamic correlations and applies a new thermal model for the temperature field in the material. The wall temperature distribution is calculated by utilising the thermal impulse and step responses of porous media, increasing the calculation speed significantly. The simulation is validated with experimental data of different wind tunnel tests and the SHEFEX II flight, showing good agreement between model and experiment. The performance of transpiration cooling is assessed by calculating the required coolant mass for different steady state and transient flight scenarios. Carbon/Carbon composite ceramic and the ultra high temperature ceramic Zirconium di-Boride are considered as wall materials. In continuous hypersonic cruise, transpiration cooling is highly effective for flight conditions with velocities below 8kms~(-1) and altitudes above 40 km. For transient re-entry, transpiration cooling is most viable for trajectories of entry velocities below 8.5 kms~(-1) and ballistic coefficients below 2.1kgm~(-2).
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摘要 :
This paper presents results of a system study of transpiration cooled thermal protection systems for Earth re-entry. The cooling performance for sustained hypersonic flight and transient re-entry of a blunt cone geometry is assess...
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This paper presents results of a system study of transpiration cooled thermal protection systems for Earth re-entry. The cooling performance for sustained hypersonic flight and transient re-entry of a blunt cone geometry is assessed. A numerical model is developed which employs semi-empirical aerodynamic correlations and applies a new thermal model for the temperature field in the material. The wall temperature distribution is calculated by utilising the thermal impulse and step responses of porous media, increasing the calculation speed significantly. The simulation is validated with experimental data of different wind tunnel tests and the SHEFEX II flight, showing good agreement between model and experiment. The performance of transpiration cooling is assessed by calculating the required coolant mass for different steady state and transient flight scenarios. Carbon/Carbon composite ceramic and the ultra high temperature ceramic Zirconium di-Boride are considered as wall materials. In continuous hypersonic cruise, transpiration cooling is highly effective for flight conditions with velocities below 8kms~(-1) and altitudes above 40 km. For transient re-entry, transpiration cooling is most viable for trajectories of entry velocities below 8.5 kms~(-1) and ballistic coefficients below 2.1kgm~(-2).
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Fully glaciated ice crystals can be ingested into aero engines, partially melt through the first few stages of compressors and eventually cause large accretions on stationary components. Ice crystals pose a threat due to damage ca...
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Fully glaciated ice crystals can be ingested into aero engines, partially melt through the first few stages of compressors and eventually cause large accretions on stationary components. Ice crystals pose a threat due to damage caused by ice shedding. With regulators expanding certification of engines to include the threat of ice crystals, there has been significant research at both the fundamental and complete engine tests. This paper details experiments which lay somewhere between the these two ends of the spectrum; an engine representative stationary geometry with direct control and measurement of the inlet icing conditions. The aim of the experiments is to directly measure ice thickness on complex three dimensional surfaces of a combined linear cascade and swan neck duct. This will enhance our understanding of locations at which there is a large threat to ice accretion and at what conditions this occurs. This paper will detail the test piece geometry, present the results of the experimental campaign and initial analysis and conclusions from the experiments.
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摘要 :
Fully glaciated ice crystals can be ingested into aero engines, partially melt through the first few stages of compressors and eventually cause large accretions on stationary components. Ice crystals pose a threat due to damage ca...
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Fully glaciated ice crystals can be ingested into aero engines, partially melt through the first few stages of compressors and eventually cause large accretions on stationary components. Ice crystals pose a threat due to damage caused by ice shedding. With regulators expanding certification of engines to include the threat of ice crystals, there has been significant research at both the fundamental and complete engine tests. This paper details experiments which lay somewhere between the these two ends of the spectrum; an engine representative stationary geometry with direct control and measurement of the inlet icing conditions. The aim of the experiments is to directly measure ice thickness on complex three dimensional surfaces of a combined linear cascade and swan neck duct. This will enhance our understanding of locations at which there is a large threat to ice accretion and at what conditions this occurs. This paper will detail the test piece geometry, present the results of the experimental campaign and initial analysis and conclusions from the experiments.
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Expansion tubes are the only type of ground test facility currently able to simulate high Mach number scramjet test flows. These access-to-space flow conditions are characterised by total pressures of the order of gigapascals. The...
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Expansion tubes are the only type of ground test facility currently able to simulate high Mach number scramjet test flows. These access-to-space flow conditions are characterised by total pressures of the order of gigapascals. The University of Queensland's X2 expansion tube facility has recently been used to generate scramjet flow conditions between Mach 10-14, with total pressures up to 10 GPa. Flow conditions were relevant to a 96 kPa dynamic pressure ascent trajectory. For ground testing of sub-scale scramjet-powered vehicles, pressure-length (p-L) scaling is used in order to maintain similarity for various flight parameters, such as Reynolds number, total enthalpy, and binary reaction rates. X2 was configured to achieve test flow static pressures considerably higher than the true flight values, while maintaining true flight velocities and temperatures, thereby demonstrating significant potential for p-L scaling, which is typically necessary since most model scramjet engines need to be tested at sub-scale. This paper details the combined analytical and numerical process used to develop new flow conditions in X2. Experimental results are presented for four new flow conditions, and axisymmetric CFD analysis is used to fully characterise test flow properties.
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摘要 :
Expansion tubes are the only type of ground test facility currently able to simulate high Mach number scramjet test flows. These access-to-space flow conditions are characterised by total pressures of the order of gigapascals. The...
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Expansion tubes are the only type of ground test facility currently able to simulate high Mach number scramjet test flows. These access-to-space flow conditions are characterised by total pressures of the order of gigapascals. The University of Queensland's X2 expansion tube facility has recently been used to generate scramjet flow conditions between Mach 10-14, with total pressures up to 10 GPa. Flow conditions were relevant to a 96 kPa dynamic pressure ascent trajectory. For ground testing of sub-scale scramjet-powered vehicles, pressure-length (p-L) scaling is used in order to maintain similarity for various flight parameters, such as Reynolds number, total enthalpy, and binary reaction rates. X2 was configured to achieve test flow static pressures considerably higher than the true flight values, while maintaining true flight velocities and temperatures, thereby demonstrating significant potential for p-L scaling, which is typically necessary since most model scramjet engines need to be tested at sub-scale. This paper details the combined analytical and numerical process used to develop new flow conditions in X2. Experimental results are presented for four new flow conditions, and axisymmetric CFD analysis is used to fully characterise test flow properties.
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摘要 :
Expansion tubes are the only type of ground test facility currently able to simulate high Mach number scramjet test flows. These access-to-space flow conditions are characterised by total pressures of the order of gigapascals. The...
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Expansion tubes are the only type of ground test facility currently able to simulate high Mach number scramjet test flows. These access-to-space flow conditions are characterised by total pressures of the order of gigapascals. The University of Queensland's X2 expansion tube facility has recently been used to generate scramjet flow conditions between Mach 10-14, with total pressures up to 10 GPa. Flow conditions were relevant to a 96 kPa dynamic pressure ascent trajectory. For ground testing of sub-scale scramjet-powered vehicles, pressure-length (p-L) scaling is used in order to maintain similarity for various flight parameters, such as Reynolds number, total enthalpy, and binary reaction rates. X2 was configured to achieve test flow static pressures considerably higher than the true flight values, while maintaining true flight velocities and temperatures, thereby demonstrating significant potential for p-L scaling, which is typically necessary since most model scramjet engines need to be tested at sub-scale. This paper details the combined analytical and numerical process used to develop new flow conditions in X2. Experimental results are presented for four new flow conditions, and axisymmetric CFD analysis is used to fully characterise test flow properties.
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This paper experimentally examines the internal and external flow characteristics of porous zirconium diboride (ZrB_2), an Ultra-High-Temperature-Ceramic (UHTC) and a potential candidate for transpiration cooling of hypersonic veh...
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This paper experimentally examines the internal and external flow characteristics of porous zirconium diboride (ZrB_2), an Ultra-High-Temperature-Ceramic (UHTC) and a potential candidate for transpiration cooling of hypersonic vehicles. This is performed for both partially sintered material and fully densified material with cast features. The Darcy and Forchheimer permeability coefficients of these samples are determined using an ISO standard test rig. The outflow of the transpiring porous samples is investigated where no hypersonic cross-flow is involved using hot-wire anemometry and focused Schlieren visualisation. The velocity maps obtained from the hot-wire data show significant non-uniformities across the UHTC's outflow region, both at low and high differential pressures. The focused Schlieren using carbon dioxide as the injected gas reveals unsteady structures at high differential pressures as the outflowing gas interacts with the surrounding air.
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